Solid propellant canister loaded multiple pulsed or staged rocket motor

ABSTRACT

A multi-pulse or multi-stage canister loaded solid propellant rocket motor. The canisters are prepared separately and loaded with solid propellant whereby the scrap rate may be reduced after which they are installed in a monolithic case which affords stiffness continuity over the length of the rocket motor to prevent guidance system upsetting free play. In order to reduce the complexity of installation of the membrane seal assembly for each pulse, the bulkhead therefor is manufactured integral with the case of the respective canister and becomes the forward closure thereof. For the multi-stage rocket motor, the monolithic case may be stepped and tapered, and the stages severable therefrom as they burn out by primer cord.

The present invention relates to solid propellant rocket motors.

For those rocket motors which have a single mass or grain of solidpropellant, the entire propulsive capacity is usually spent during thecombustion process thereof. This is for the reason that once a solidpropellant grain is ignited it is very difficult to stop the combustionprocess until the entire mass of ignited propellant has been consumed.

In order to provide a "start-stop restart" capability so that a rocketmotor will have an ability to fire more than once for greatermanuevering flexibility, solid propellant rocket motors have beenprovided with multiple pulses, such as described in U.S. Pat. No.4,766,726 to Tackett et al, assigned to the assignee of the presentapplication, and multiple stages. The aforesaid U.S. Pat. No. 4,766,726is incorporated herein by reference. A multiple pulsed rocket motor isone which contains two or more solid propellant grains such as a boostgrain and a sustain grain which are separated by a membrane sealassembly or other means to enable the ignition of the solid propellantgrains to be independent of each other whereby discrete impulses areavailable upon command. Where the solid propellant grains are positionedin tandem with each other, that is, with one solid propellant grainforward of the other, the membrane seal assembly has extended over theinner diameter of the rocket motor case and has been attached to therocket motor case. The membrane seal assembly includes a bulkhead whichincludes a plurality of apertures for flow of combustion gasestherethrough and also includes a thin imperforate metallic membrane orcover of high strength but ductile material which covers the aft side ofthe bulkhead to seal the forward chamber from flow of gases thereintoupon ignition of the solid propellant grain in the aft chamber andwhich, after the solid propellant grain in the aft chamber has beenexpended, pressure resulting from combustion of the solid propellantgrain in the forward chamber upon ignition thereof at a selected timewill cause the thin membrane to rupture and thus allow the escape ofgases from the forward chamber through the apertures in the bulkhead tothe aft chamber and then out the nozzle to produce thrust.

A multiple staged rocket motor also provides a "start-stop-restart"capability. Like the multiple pulsed rocket motor, each stage has aseparate solid propellant grain. However, a separate nozzle is providedfor each grain and the grain in each stage is separated from the grainsof other stages so that, when a grain has been consumed, the stagecontaining that grain is caused by explosive bolt means or the like toseparate from the adjoining stage to thus remove excess weight from therocket motor so that increased range and/or speed may be achieved. Theadjoining stage with its separate grain and nozzle may then be fired ata selected later time during the flight of the rocket.

Pulsed and staged rocket motors have typically been provided withsegmented cases to allow for separate processing of the boost andsustain propellant grains.

When a pulsed or staged segmented rocket motor is coasting and the jointis not loaded, such as during interpulse delay, the segments may have atendency to move relative to each other. Free play or movement withinthe tolerance space of the connectors can undesirably upset a missile'sguidance system and thus throw a missile off course.

Segmented multi-stage or multi-pulse solid propellant rocket motors arecommonly pinned or bolted together with complex attachment areas whichare heavy, costly to process, and may undesirably allow leaks betweenthe stages or pulses. Furthermore, it may be difficult to reliablyattach separate bulkheads to the segmented multi-pulse motor case.

It is, therefore, an object of the present invention to provide amultiple pulsed or staged solid propellant rocket motor wherein there isadequate stiffness continuity along the length of the case to insurethat the missile's guidance system is not upset.

It is another object of the present invention to provide such a rocketmotor which is rugged and reliable yet easy to manufacture and whichallows separate processing of the propellant grains.

It is a further object of the present invention to provide such a rocketmotor wherein the bulkheads are more easily and reliably providedbetween pulses.

These and other objects of the invention will become apparent in thefollowing detailed description of the preferred embodiment of theinvention taken in connection with the accompanying drawings.

IN THE DRAWINGS

FIG. 1 is a sectional longitudinal view of a pulsed rocket motor whichembodies the present invention;

FIGS. 1A, 1B, 1C, and 1D are detail views of portions A, B, C, and Drespectively of the rocket motor of FIG. 1;

FIG. 2 is a detail view of a portion of the forward closure of therocket motor of FIG. 1;

FIG. 3 is a detail view of a portion of a bulkhead pulse separator forthe rocket motor of FIG. 1;

FIG. 4 is a detailed view of a portion of the aft closure of the rocketmotor of FIG. 1;

FIGS. 5 to 13 are schematic views illustrating steps in the manufactureof the rocket motor of FIG. 1;

FIGS. 14 to 16 are schematic longitudinal views which illustrate insequence the operation of the rocket motor of FIG. 1; and

FIG. 17 is a schematic longitudinal view of a multiple staged rocketmotor which embodies the present invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Referring to FIG. 1, there is shown generally at 10 a solid propellantrocket motor which has three solid propellant combustion chambers orpulses 12, 14, and 16. Greater details of portions of the rocket motorare shown in FIGS. 1A, 1B, 1C, 1D, and 2 to 4. For example, whereas FIG.1, for ease of illustration, does not show insulation, it is clearlyshown in FIGS. 2 to 4. In order to provide stiffness continuity alongthe length of the rocket motor 10 in accordance with the presentinvention so that the missile's guidance system may more accuratelymaintain the missile on course, the rocket motor 10 is provided with amonolithic elongate generally cylindrical tube or case 18 which extendsgenerally over the entire length of the rocket motor 10. The case 18 maybe composed of any suitable material such as a metallic or compositematerial which provides adequate strength and stiffness. For example,the case 18 may be a filament wound composite of carbon or graphitefibers, such as carbon fibers marketed by Toray Industries Inc. ofTokyo, Japan, under the designation T-40, impregnated with a polyimideresin, such as that commonly known as PMR-15, for high temperaturecapacity. Alternatively, for greater strength, a bis-maleimid resin maybe used. The windings may, for example, be in ahelical/hoop/helical/hoop/helical pattern.

In order to allow separate manufacture and processing of the solidpropellant grain for each of the pulses for ease of manufacture of therocket motor 10 in accordance with the present invention, each of thepulses 12, 14, and 16 comprises a canister 20 which may be manufacturedand loaded separately and then installed in the monolithic case 18.

Each canister 20 comprises a thin canister wall 22, which may, forexample, be composed of stainless steel such as 15-5 PH stainless steeland in which is loaded any suitable conventional propellant material,illustrated at 24, such as, for example, that commonly known as GAPpropellant, i.e., a propellant containing an energy enhancing glycidalazide polymer binder. The propellant grains 24 may be centrallyperforated as shown by perforations, illustrated at 40 in FIGS. 2, 3,and 4, which may be configured in any suitable conventionalconfiguration such as star-shaped or have radial slots in accordancewith the particular requirements of the rocket motor. As shown in FIGS.2, 3, and 4, any suitable conventional insulation material 64, such asan aramid fiber impregnated rubber, is provided between the grain 24 andcanister wall 22. Stress relieving flaps 66 are suitably provided byslits 68 in the insulation 64 in accordance with principles known tothose of ordinary skill in the art to which this invention pertains. Asuitable conventional liner (not shown) may be provided between thegrain 24 and insulator 64.

The canisters 20 are inserted into the monolithic case 18 in end-to-endrelation and are attached to each other as hereinafter described. Eachof the canisters 20 is attached or otherwise suitably bonded to themonolithic case 18 by a suitable bonding material 26 such as, forexample, an epoxy marketed by Armstrong World Industries, Inc. ofLancaster, Pa., under the designation A661 less glass filler material,having a thickness of perhaps 0.03 inch. The cure temperature of thebonding material 26 should be lower than the sensitivity of thepropellant and sufficiently low so as not to degrade the bondlines ofthe propellant to the insulation, i.e., the cure temperature ispreferably not over about 200° F.

Each pair of canisters is separated by a domed closure member orbulkhead illustrated at 28, which may for example be composed ofstainless steel such as 15-5 PH stainless steel which will be describedin greater detail hereinafter. The bulkhead is part of a membrane sealassembly 30. The membrane seal assembly 30 includes a through bulkheadfitting or boss 32 (see FIG. 3) in which is sealingly inserted anassembly for positioning a suitable igniter 36 such as, for example, aconsumable pyrogen igniter aft of the bulkhead 28 for igniting the grain24 in the canister 20 associated with the membrane seal assembly 30,i.e., the grain aft of the membrane seal assembly. A suitable initiator34, which is connected via connector 39 to a source (not shown) ofelectrical energy by suitable leads 38, such as fiber optic leads, whichare routed thereto from forwardly of the rocket motor 10 through thecentral apertures or perforations 40 in the grains 24 which are in thecanister or canisters 20 forward thereof in a conventional manner asshown. The igniter assembly is sealed in the feed-through boss 32 by asuitable seal such as a glass to metal seal or an epoxy to preventescape of gases therethrough. Such a seal may be provided usingprinciples of common knowledge to those of ordinary skill in the art towhich this invention pertains, and may be provided similarly asdescribed in the aforesaid U.S. Pat. No. 4,766,726.

Referring to FIG. 2, there is shown a detailed view of a portion of theforward closure. A generally cylindrical stub skirt 42 composed of asuitable material such as 15-5 PH stainless steel and having a thicknessof perhaps 0.03 inch is inserted in the upper or forward end portion ofthe case 18 is bonded thereto by bonding material 26 to increase bondarea for reacting the dome loads into the case 18. The stub skirt 42 hasan enlarged forward portion 44 which contains a circumferential groove46 in its radially outer surface which faces the case 18. In the groove46 may be contained an O-ring 48 to provide a wiper seal to provide avacuum to draw in the bonding material 26 as the assembly of canistersare loaded into the case 18. Alternately, the bonding material may beinjected between the case 18 and canisters 20 under pressure in whichcase the wiper seal is not needed. The aft end portion 50 of the stubskirt is enlarged, and the enlarged portion 50 is threaded.

The canister 20 for the forward or third pulse 16 is composed ofstainless steel such as 15-5 PH stainless steel or other suitablematerial and has the aforesaid generally cylindrical wall 22 with athickness of perhaps 0.03 inch which is integral with a forward domeportion 52 which has an enlarged thickness of perhaps 0.090 inch towithstand longitudinal loads to form the forward closure. at theintersection of the generally cylindrical wall 22 and the dome 52 is anenlarged portion 56 which is threaded to mate with threads on thethreaded portion 50 of the stub skirt 42 to provide a threaded jointillustrated at 58 between the canister 20 for the third pulse 16 and thestub skirt 42. The enlarged portion 56 also has a groove 60 in aradially outer surface thereof facing the stub skirt aft portion 50 inwhich is inserted an O-ring 62 to provide a seal against escape of gasesforwardly of the the third pulse 16. Centrally of the dome 52 is afeed-through boss 32, as previously described, for sealingly insertingthe initiator/igniter assembly for the third pulse 16.

A more detailed view of a portion of the membrane seal assembly 30between two pulses, i.e., pulses 14 and 16, is shown in FIG. 3, themembrane seal assembly between pulses 12 and 14 being similar. For thepurposes of description of FIG. 3, pulse 16 will be called the forwardpulse and pulse 14 will be called the aft pulse. The canister 20 for theforward pulse 16 terminates at its afte end in an enlarged end portion70 which is threaded similarly to the end portion 50 of the stub skirt42. The canister 20 for the aft pulse 14, which is also composed of 15-5PH stainless steel or other suitable material, has the wall 22 aspreviously described with a thickness of perhaps 0.03 inch and a domeshaped bulkhead 28 integral therewith which has a greater thickness ofperhaps 0.09 inch to withstand longitudinal loads to form a closurebetween the forward and aft pulses 16 and 14 respectively. If a canister20 for a pulse, other than the forward pulse, is composed of a compositematerial, the bulkhead 28 therefor may comprise thin stainless steelbonded inside of the domed portion of the canister for attachment of themembrane 82. A canister is preferably composed of thin stainless steelor other suitable metal in order that the assembly of the membrane sealassembly 30 may be easier.

At the intersection of the wall 22 and bulkhead 28 is an enlargedportion 72, similar to enlarged portion 56 in FIG. 2, which is threadedto mate with threads on the forward canister aft portion 70 to provide athreaded joint illustrated at 74 therebetween. Enlarged portion 72includes a groove 76 in its radially outer surface facing the canisterend portion 70 in which is inserted an O-ring 78 to seal between thesecond and third pulses 14 and 16 respectively.

In order to provide a flow of gases through the bulkhead 28 into thecanister 20 for the aft pulse 14 so that they may be expelled throughthe nozzle during burning of the propellant grain 24 in the canister 20for the forward pulse 16, a plurality of apertures 80 are contained inthe bulkhead 28. Apertures 80 are preferably of a sufficient size andquantity to provide a flow area therethrough which is equal toapproximately four times the nozzle area so as to minimize pressure droptherethrough but which are also preferably positioned to provide maximummaterial strength in accordance with principles commonly known to thoseof ordinary skill in the art to which this invention pertains. A thinmetallic burst diaphragm 82 composed of any suitable material such as,for example, stainless steel or nickel is provided along the aft surfaceof the bulkhead 28 and is bonded thereto by any suitable means such as,for example, welds illustrated at 84 to prevent the flow of gasesthrough the apertures 80 while the propellant grain 24 in the canister20 for the aft pulse 14 is burning but which is caused to rupture toallow the passage of gases through the apertures 80 after the aft pulsegrain 24 has completed burning and during burning of the propellantgrain in the canister 20 for the forward pulse 16. In other words, thediaphragm 82 is caused to rupture to provide flow communication throughthe apertures 80 between a pair of adjacent canisters 20 when a higherpressure is present in the forward one of the canisters than in theafter one of the canisters. The burst diaphragm 82 may be scored asdiscussed in the aforesaid patent application which is to issue as U.S.Pat. No. 4,766,726 to rupture easily and predictably upon a pressure inthe forward canister 16 exceeding the pressure in the aft canister 14 bya substantial amount. Centrally of the bulkhead 28 is a feed-throughboss 32 integral therewith to permit insertion and positioning of theigniter assembly for ignition of the respective propellant grain 24 aftof the bulkhead 28, as previously discussed with respect to the forwardclosure of FIG. 2. Insulation 64 is provided between the grain 24 andthe bulkhead 28 as well as between the grain 24 and the wall 22 for thecanister 20 for the aft pulse 14, the insulation next to the bulkheadbeing provided with a stress relieving flap 66. In the canister 20 forthe forward pulse 16, insulation 64 is provided to cover the bulkhead 28leaving the apertures 80 uncovered, and the aft end portion of theinsulation along the cannister wall 22 is provided with a flap 66, asshown in FIG. 3.

Referring to FIG. 4, there is shown a portion of the aft closure of therocket motor 10. The canister 20 for the first pulse 12 extendsaftwardly to terminate, similarly to the canister for the third pulse16, in an enlarged threaded portion 86. An aft closure member 88 has anenlarged threaded portion 90 to threadedly mate with the canisterthreaded portion 86 to provide a threaded joint illustrated at 92therebetween. A groove 94 is provided in the radially outer surface ofthe enlarged portion 90 facing the canister portion 86, and an O-ring 96is provided in the groove 94 to provide a seal against escape of gasesbetween the aft closure member 88 and the canister 20. A generallycylindrical portion 98 of the aft closure member 88 extends rearwardlyfrom the enlarged portion 90 and is suitably bonded with bondingmaterial 26 to the case 18. A dome shaped portion 100 of the aft closuremember 88 extends rearwardly and inwardly from the enlarged portion 90to provide means for attachment of nozzle 102 in accordance withprinciples commonly known to those of ordinary skill in the art to whichthis invention pertains. The nozzle throat may be composed of tungstenor of rhenium coated carbone graphite to reduce throat erosion forconsistent performance. Insulation 64, which may be provided with asuitable relief flap 66 in the area of the threaded joint 92, is causedto cover the inner surfaces of the canister wall 22, the aft closuremember 88, and the axially inner surface of the nozzle 102.

Each canister 20 is manufactured and loaded with a grain 24 separately.The loaded canisters 20 are then assembled together by means of thethreaded joints 74, as previously discussed, or other suitable means.The aft closure member 88 is attached to the aft canister 20 by means ofthreaded joint 92, as previously discussed, or other suitable means. Thestub skirt 42 is attached to the forward canister 20 by means of thethreaded joint 58 as previously discussed, or by other suitable means.This assembly is then inserted in and bonded into the monolithic case 18which is the primary load carrying structure.

FIGS. 5 to 13 illustrate in sequence the manufacture of a rocket motoraccording to the present invention. After the monolithic case 18 hasbeen filament wound by a suitable filament winding process, asillustrated in FIG. 5, it is suitably cured at an elevated temperatureof perhaps 600° F. at 150 psi for a period of perhaps 9 hours in anautoclave, as illustrated in FIG. 6. The mandrel is then removed, andthe case hydroproofed and any domes removed, as illustrated in FIG. 7.In the case 18 is produced by a continuous braider, then the case 18, ofthe desired length, is cut therefrom. The canisters 20 are fabricated,as illustrated in FIG. 8, either before, after, or while the case 18 ismade, loaded with solid propellant grains 24 and inspected by use ofsuitable inspection apparatus 112, as illustrated in FIG. 9, and thenassembled together, as illustrated in FIG. 10 and as previouslydiscussed. As also previously discussed, the stub skirt 42 and aftclosure member 88 are also attached. The assembled canisters are theninserted in the monolithic case 18 and bonded thereto, as illustrated inFIG. 11, by pumping the low cure temperature bonding material 26 betweenthe canister walls 22 and the case 18 under pressure and curing it, asillustrated in FIG. 12. The bond may be inspected by suitable inspectionapparatus 114 as illustrated in FIG. 13 after which the bond may beproof tested with inert gas. The nozzle 102 may then be installed.

For example, the monolithic case 18 may have a length of about 82inches, a diameter of about 8 inches, and a thickness of about 0.12 inchand have three pulses the canister 20 for each having a length of about23 inches and a thickness of about 0.03 inch. The case 18 may, forexample, be a filament wound composite of resin impregnated carbon orgraphite fibers, as previously discussed, while each of the canistersmay, for example, be composed of 15-5 PH stainless steel and be loadedwith a conventional propellant material including a conventionalinsulator and liner, as previously discussed. The nozzle throat may, forexample, have a diameter of about 1.34 inches.

FIGS. 14 to 16 illustrate in sequence the operation of rocket motor 10of FIG. 1. The arrows 106 illustrate internal pressure due to combustionof the propellant material 24. The arrows 108 illustrate the flow ofgases generated from burning of the propellant grains 24. The arrows 110ilustrate the points where the longitudinal load from closures reactthrough the bond line into the case 18. The rocket motor 10 is initiallyoperated by providing power to the initiator 34 through leads 38 toignite the igniter 36 for the first pulse 12 to ignite the solidpropellant grain 24 in the canister 20 for the first pulse 12, asillustrated in FIG. 14. The internal pressure 106 is contained withinthe canister 20 for the first pulse 12 and the gases which are producedare released through the nozzle 102 to produce thrust for forwardmovement of the rocket motor. The bulkhead 28, with the apertures 80therein covered by the burst diaphragm 82, prevent the passage of gasesinto the canisters 20 for the second and third pulses 14 and 16respectively.

After the first pulse has been extinguished, the solid propellant grain24 for the second pulse 14 may be ignited at a selected time. When it isignited, the pressure of the gases generated therefrom burst thediaphragm 82 between the first and second pulses 12 and 14 respectivelyand pass through the apertures 80 thereof into the canister 20 for thefirst pulse 12 and are then released through the nozzle 102 to producethrust, as illustrated in FIG. 15. The bulkhead 28 and burst diaphragm82 which separate the second and third pulses 14 and 16 respectivelyprevent the passage of the generated gases into the canister 20 for thethird pulse 16.

As illustrated in FIG. 16, the bulkhead 28 and burst diaphragm 82between the first and second pulses 12 and 14 respectively aresubstantially consumed during the second pulse. After the second pulsehas been extinguished, the solid propellant grain 24 for the third pulse16 may be ignited at a selected time. The pressure generated by thegases from burning the grain 24 for the third pulse 16 burst thediaphragm 82 between the second and third pulses 14 and 16 respectivelyand pass through the apertures 80 in the bulkhead therebetween afterwhich they are passed through the first and second pulse canisters thenthrough the nozzle 102 to provide thrust.

Referring to FIG. 17, there is schematically shown generally at 200 arocket motor in accordance with an alternative embodiment of the presentinvention. The rocket motor 200 comprises a monolithic case 202, whichmay be composed of any suitable material such as stainless steel orcarbon or graphite fiber and resin composite material similarly asdescribed for the case 18 of FIG. 1, which is closed at its forward end204 such as by a domed structure 206 and is open at its aft end 208.Contained within the monolithic case 202 in end-to-end relation is aplurality of stages 210, 212, and 214 each of which comprises a canister216 loaded with a suitable conventional solid propellant illustrated at218 and a nozzle 220 attached to the canister 216 at its aft end forexpelling gases from combustion of the propellant 218 for providingthrust.

Each of the canisters 216 has a generally dome-shaped closed integralforward end portion 222 which is suitably reinforced such as byincreased wall thickness to prevent escape of population gases forwardthereof to other stages.

The case 224 of each canister may be composed of any suitable materialsuch as an insulation material overwrapped with carbon fiber/epoxycomposite. Shrink flaps have been conventionally used between the caseand insulator of conventional rocket motors to prevent cracking ordebonding of the propellant from the insulator. In order to eliminatebondline problems caused by propellant shrinkage without the need toprovide such shrink flaps between the case and insulator, it ispreferred that the canister case 224 be composed of a transversely(transverse to the fiber direction) elastomeric reinforcing materialsuch as a composite of a resin impregnated continuous fiber such asaramid oriented at a low angle such as less than 15 degrees relative tothe case axis coated with rubber or other suitable elastomeric materialand polar wound to form the canister case. The polar windings may becompacted by any suitable means such as curing in a female mold. If hoopwindings are applied for compacting, they should be removed afterwardsto allow for contraction and expansion. Other suitable fibrous materialsuch as carbon or graphite fibers or fiberglass material may alternatelybe used. If desired, the canister case 224 may be composed of thinstainless steel or other metal with insulation material appliedinternally thereof. The insulation may be a rigid rubber or otherwisesufficiently rigid for the nozzle 220 to be attached directly in amanner and using principles commonly known to those of ordinary skill inthe art to which this invention pertains. Alternatively, a polar boss(not shown) may be attached to the canister case 224, and the nozzle 220attached to the polar boss. The case 224 is preferably cured in a femalemold to provide a suitable surface of bonding into the monolithic case202. Propellant material 218 is preferably cast therein while in thesame mold to prevent its contamination with hydroscopic moisture fromthe air.

In order to allow increased grain configuration flexibility fortailoring of the burn rate pattern and to provide ease of core removalafter the propellant is cast, propellant material 218 in the canister216 may desirably be segmented and the segments suitably individuallybonded or otherwise attached in the canister case 202 in accordance withprinciples commonly known to those of ordinary skill in the art to whichthis invention pertains.

Each of the canister cases 224 is preferably bonded or otherwisesuitably attached to the monolithic case 202, with the aid of aconventional retainer (not shown) which fits in a groove of the case 202to lock the canister therein, using a room temperature cured epoxy,illustrated at 226, such as a versamide or tetro-amine cured epoxy so asnot to unduly heat the propellant. Alternatively, the canister cases 224may be pinned or screwed, such as by a modified buttress thread, intothe monolithic case 202. In order to allow wedging of the canister cases224 in the monolithic case 202 so that they are seated in place for atight fit during acceleration of the rocket motor 200 so that a case 224is not "rammed" into a case forward thereof, the portions 230 of themonolithic case 202 to which the individual canister cases 224 are notbonded, i.e., the interstage portions, are tapered so that the case 202has a smaller diameter at its forward end 204 than at its aft end 208,and the diameters of the canister cases 224 are accordingly decreasedfrom the aft canister to the forward canister. However, the portions 232of the monolithic case 202 to which the individual canister cases arebonded are stepped, i.e., not tapered, to prevent the necessity ofhaving to also taper the canister cases 224. Servicing access ports (notshown) for final connections of conventional vector control andelectrical devices may be cut through interstage portions 230 of themonolithic case 202.

In accordance with the present invention, suitable means are providedfor severing the monolithic case 202 to release each of the canisters216 in turn after the solid propellant material 218 therein is burned.Such means preferably comprises primer cord 228, that is, a fast burninglongitudinal explosive such as a putty or disc igniterant for cutting,which is preferably bonded along the inner surface of the monolithiccase 202 circumferentially thereof between each pair of canisters 216,i.e., in each of the tapered interstage portions 230. At a selectedtime, the primer cord 228 is ignited to sever the monolithic case 202 sothat the stage aft thereof is released from the rocket motor.

The monolithic case 202 and canisters 216 are constructed separatedly.After each canister case 224 has been constructed it is loaded withsolid propellant 218 and a nozzle 220 is attached as previouslydiscussed. Each of the canisters 216 is then inserted in turn in themonolithic case 202 and bonded to the respective stepped portion 232thereof. The primer cord 228 is bonded about the circumference along theinner surface of each interstage portion 230 after the respectivecanisters 216 forward thereof have been installed.

A conventional igniter illustrated schematically at 234, which mayinclude a suitable initiator, is provided for igniting the solidpropellant grain 218 in each of the canisters 216. Each igniter 234 isconnected by lead wires 236 to a source of power (not shown).

Operation of the rocket motor 200 is initiated by supplying powerthrough lead wires 236 to the igniter 234 for the canister 216 for thefirst stage 210. The igniter 234 is caused to ignite the propellant 218in the first stage 210 whereby the propellant burns to produce gaseswhich are expelled through the nozzle 220 for the first stage 210 toprovide thrust for forward movement of the rocket motor. After it hasburned out, the first stage 210 is removed from the rocket motor 200 byigniting the primer cord 228 in the interstage portion 230 between thefirst and second stages 210 and 212 respectively to sever the monolithiccase 202. At a selected time during the flight of the rocket motor 200,the second stage 212 is ignited similarly to the ignition of the firststage to produce thrust. After the second stage is burned out it issevered from the rocket motor similarly to the severing of the firststage therefrom. At another selected time during the flight, the thirdstage 214 is ignited to produce additional thrust for carrying therocket to its destination.

Alternately, the monolithic case may be only slightly tapered over itslength to permit wedging of cylindrical canisters therein. If desired,the loaded canisters may be frozen or suitably lowered in temperaturethen loaded upside down in the monolithic case so as to expand and forma tight fit therewith as they warm to ambient temperature with a resultthat it may be unnecessary to taper the monolithic case at all. Asuitable bonding agent or mechanical interlock may also be provided inaddition to the tight fit to provide positive load transfer. If themonolithic case is composed of a material such as a carbon or graphitefiber material providing a negative or only a small positive coefficientof thermal expansion, the case may be loaded with the canisters while italso at the lower temperature whereby the canisters may expand more thanthe case expands during warming to ambient temperature to form the tightfit.

The monolithic case 18 or 202 is provided to have sufficient strengthand stiffness to carry longitudinal and hoop loads due to internalpressure as well as external loads due to flight environments orhandling. Thus, the monolithic case 18 or 202 is provided to function asa stress, stiffening, and attachment complement to the pulses 12, 14,and 16 or stages 210, 212, and 214 to eliminate indesirable free playalong the length of the rocket motor and allow the canisters 20 or 216to have thinner walls as previously discussed since they need not bebuilt to assume the stresses and stiffening which are assumed by themonolithic case 18 or 202. Since only pulses or stages and not an entiremotor need be scrapped, the present invention allows reduced scrap.Since rocket motor 10 or 200 does not require complex case joint ringassemblies, cost, weight, inert volume, and required hardware may beadvantageously reduced. Furthermore, a rocket motor according to thepresent invention may require little or no "blind hole" fabrication. Themonolithic case of the present invention may be fabricated inexpensivelyand easily since it does not require any domes, polar bosses, skirts,shear plies, or joints. The monolithic case 18 is also provided tosimplify the membrane seal assembly 30 construction wherein the bulkheadthereof may be integral with the canister wall and thus more reliablyform the forward closure for each canister.

It is to be understood that the invention is by no means limited to thespecific embodiments which have been illustrated and described hereinand that various modifications thereof may indeed be made which comewithin the scope of the present invention as defined by the appendedclaims.

What is claimed is:
 1. A rocket motor comprising a monolithic elongategenerally cylindrical case having means defining an aft aperture, atleast two canisters each of which has an individual case and contains anindividual grain of solid propellant material, said canisters disposedsequentially within said monolithic case in end to end relation withsaid canister cases attached to said monolithic case and to each other,means forming a closure to preclude flow communication between saidcanisters which includes means providing flow communication between saidcanisters when a higher pressure is present in the forward one of saidcanisters than in the aft one of said canisters, means for igniting eachof said solid propellant grains, and a thrust nozzle means attached tosaid monolithic case in flow communication with said aft aperture.
 2. Arocket motor according to claim 1 wherein said closure means comprises abulkhead for providing a forward closure of the aft one of saidcanisters, said bulkhead including a plurality of perforation means forproviding flow communication between said canisters, and a rupturablemembrane seal means covering said perforation means to preclude flowcommunication between said canisters and to rupture and thereby allowflow communication between said canisters when a higher pressure ispresent in the forward one of said canisters then in the aft one of saidcanisters.
 3. A rocket motor according to claim 2 wherein said bulkheadis integral with said aft canister case.
 4. A rocket motor according toclaim 1 wherein said individual canister case are bonded to saidmonolithic case.
 5. A method of making a rocket motor comprising thesteps of:a. providing a monolithic elongate generally cylindrical casehaving an aft aperture; b. providing at least two canisters each havingan individual case and containing an individual grain of solidpropellant material; c. providing a closure on an end portion of one ofthe canisters to be attached to an open end portion of the other of thecanisters to preclude flow communication between the canisters when theyare connected together in end to end relation; d. providing means forproviding flow communication between the canisters when a higherpressure is present in said other canister then in said one canister; e.connecting said end portion of said one of the canisters to said openend portion of said other of the canisters; f. inserting the connectedcanisters in the monolithic case; g. attaching the individual canistercases of the connected canisters to the monolithic case; and h.attaching a thrust nozzle to the monolithic case in flow communicationwith the aft aperture.
 6. A method according to claim 5 wherein the stepof providing a closure comprises providing a bulkhead which closes saidend portion of said one of the canisters, the step of providing flowcommunication means comprises providing a plurality of perforations inthe bulkhead and covering the perforations with a membrane whichprecludes flow communication between the canisters and which rupturesand thereby allows flow communication between the canisters when ahigher pressure is present in said other of the canisters than in saidone of the canisters.
 7. A method according to claim 6 furthercomprising forming the bulkhead integral with the case of said one ofthe canisters.
 8. A method according to claim 5 wherein the step ofattaching said individual canister cases to the monolithic casecomprises bonding said individual canister cases to the monolithic case.